A methodology for the design of small autonomous underwater vehicle propulsion systems has been developed and applied to the Virginia Tech 690 AUV. The methodology is novel in that it incorporates fast design level codes capable of predicting the viscous effects of low Reynolds number flow that is experienced by small, slow turning propellers. The methodology consists of determining the minimum induced loss lift distribution for the propeller via lifting line theory, efficient airfoil sections for the propeller via a coupled viscous-inviscid flow solver and optimization, brushless DC motor identification via ideal motor theory and total system efficiency estimates. The coupled viscous-inviscid flow solver showed low Reynolds number flow effects to be of critical importance in the propeller design. The original Virginia Tech 690 AUV propulsion system was analyzed yielding an experimental efficiency of 26.5%. A new propeller was designed based on low Reynolds number airfoil section data yielding an experimental efficiency of 42.7%. Finally, an entirely new propulsion system was designed using the methodology developed herein yielding a predicted efficiency of 57-60%. / Master of Science

The original intention regarding the Reynolds’ Number Apparatus was to design, construct, and operate extensively the experimental machinery. The extensive operation of the apparatus, however, would entail not only the investigation of the effects of various conditions on the Reynolds’ number of liquids, but also the effects of laminar and turbulent flow on the heat transferred from the liquid to the pipe walls. It would also be possible to use the apparatus to determine the friction factors of various diameter pipes. Any one of the above subjects contains enough material for a thesis. It was therefore decided to limit the scope of this investigation to the design, construction, and calibration of the apparatus. In view of the extremely slow delivery of equipment and the many delays encountered during the construction, this was indeed a wise decision. It may therefore be stated that the purpose of this report is to present to the reader the Reynolds’ Number Apparatus and to acquaint him with its design, construction, calibration, and experimental possibilities. Although only enough test runs will be made to calibrate the instruments required to determine the Reynolds’ number of a given oil at one temperature, it is the author’s belief that this report will serve as an adequate guide to the construction of similar apparatuses and will stimulate others to a more extensive investigation of the experimental possibilities. / Master of Science

The presence of roughness on a surface subject to high Reynolds number flows promotes the formation of a turbulent boundary layer and the generation of a fluctuating pressure field imposed on the surface. While numerous studies have investigated the wall pressure fluctuations over zero-pressure gradient smooth walls, few studies have examined the effects of surface roughness on the wall pressure field. Additionally, due to the difficulties in obtaining high Reynolds number flows over fully rough surfaces in laboratory settings, an even fewer number of studies have investigated this phenomenon under flow conditions predicted to be fully free of transitional effects that would ensure similarity laws could be observed. This study presents the efforts to scale and describe the wall pressure spectrum of a rough wall, high Reynolds number turbulent boundary layer free of transitional effects. Measurements were taken in the Virginia Tech Stability Wind Tunnel for both smooth and rough walls. A deterministic roughness fetch composed of 3-mm hemispheres arranged in a 16.5-mm square array was used for the rough surface. Smooth and rough wall flows were examined achieving Reynolds numbers up to Re_θ = 68700 and Re_θ = 80200 respectively, with the rough wall flows reaching roughness based Reynolds numbers up to k_g⁺ = 507 with a simultaneous blockage ratio of δ/k_g = 76. A new roughness based inner variable scaling is proposed that provides a much more complete collapse of the rough wall pressure spectra than previous scales had provided over a large range of Reynolds numbers and roughness configurations. This scaling implies the presence of two separate time scales associated with the near wall turbulence structure generation. A clearly defined overlap region was observed for the rough wall surface pressure spectra displaying a frequency dependence of Ï ^-1.33, believed to be a function of the surface roughness configuration and its associated transport of turbulent energy. The rough wall pressure spectra were shown to decay more rapidly, but based on the same function as what defined the smooth wall decay. / Master of Science

The critical Reynolds Number in this thesis is determined by the method of pipe function. This method is based on the Poiseuille's law of laminar flow. The film coefficients of the SAE 20 oil are obtained from the equation, h = Q/AΔT Where h = film coefficient of oil, in BTU/(sq ft) (hr) (°F) Q = rate of heat flow, in BTU/hr A = total cooling surface, in sq ft ΔT = mean temperature difference between the oil and the pipe wall, in °F The rate of heat flow is calculated from the temperature drop and the rate of flow of oil. The cooling surface is obtained by multiplying the actual inside periphery of 3/8 pipe by the total length of the heat exchanger. The mean temperature difference is solved by the method of balance of energy. The critical Reynolds Numbers obtained lie between 1700 and 2360. The pump pressure causes the vibration of the pipes and the initial turbulence of flow, and consequently has the most dominating effect on the critical Reynolds Number. The temperature and size of pipe effect the pump pressure required for testing reaching the transition region but have little direct effect on the critical value. The pressure drop for laminar flow is approximately proportional to the Reynolds number. The film coefficients of laminar flow are very low and approximately proportional to Re⁰⋅⁴. However, when the transition region is reached the film coefficients increase suddenly and rapidly, and more and more slowly as the Reynolds Number is further increased. For the same Reynolds Number, the hotter oil has the lower film coefficient. / M.S.

In order to protect gas turbine blades from hot gases exiting the combustor, several intricate external and internal cooling concepts are employed. High pressure stage gas turbine blades feature serpentine passages where rib turbulators are installed to enhance heat transfer between the relatively colder air bled off from the compressor and the hot internal walls. Most of the prior studies have been restricted to Reynolds number of 90000 and several studies have been carried out to determine geometrically optimized parameters for achieving high levels of heat transfer in this range of Reynolds number. However, for land-based power generation gas turbines, the Reynolds numbers are significantly high and vary between 105 and 106. Present study is targeted towards these high Reynolds numbers where traditional rib turbulator shapes and prescribed optimum geometrical parameters have been investigated experimentally. A steady-state liquid crystal thermography technique is employed for measurement of detailed heat transfer coefficient. Five different rib configurations, viz., 45 deg., V-shaped, inverse V-shaped, W-shaped and M-shaped have been investigated for Reynolds numbers ranging from 150,000 to 400,000. The ribs were installed on two opposite walls of a straight duct with aspect ratio of unity. For very high Reynolds numbers, the heat transfer enhancement levels for different rib shapes varied between 1.3 and 1.7 and the thermal hydraulic performance was found to be less than unity. / Master of Science / Gas turbine blades operate in hot gases exiting from combustor. The temperature of the hot gas is much higher than the melting point of blades material. To protect gas turbine blades several intricate external and internal cooling technique have been applied. Inside the blades, impingement cooling, rib turbulators cooling and pin fins cooling technique are applied in the leading edge, central body and trailing edge, respectively. At the central body serpentine passage was manufactured where rib turbulators are installed to enhance heat transfer between the relatively colder air bled off from the compressor and the hot internal walls. This is attributed to the colder air’s boundary layer is tripped by the rib turbulators enhance the flow turbulence. All the previous works are based on lower Reynolds number (under 90000) which always happens in aircraft gas turbine engine. In land based gas turbine the Reynolds numbers of cooling air are significantly high and vary between 10⁵ and 10⁶ . Present study is targeted towards these high Reynolds numbers where traditional rib turbulator shapes and prescribed optimum geometrical parameters have been investigated experimentally. Five different rib configurations, viz., 45 deg., V-shaped, inverse V-shaped, W-shaped and M-shaped have been investigated for Reynolds numbers ranging from 150,000 to 400,000. For very high Reynolds numbers, the heat transfer enhancement levels for different rib shapes varied between 1.3 and 1.7 and the thermal hydraulic performance was found to be less than unity. It’s a caution to turbine hot gas path designers, particularly for the cases where rib designs for aircrafts are used in land based power generationgas turbines

In this study the primary objective was to determine the effect of aspect ratio in particular and in general the effect of three dimensionality on the flow around wings at low Reynolds numbers. It was seen that the effects observed at high Reynolds number are also present in this Re range. There is the usual increase in lift slope and this increase can even be predicted with reasonable accuracy using Prandtl's lifting line theory. In addition to the change in lift slope the zero lift angle of attack was also influenced by the aspect ratio. Through flow visualization it was ascertained that the wingtips have a rather restricted effect on the laminar separation bubble. The disappearance of the bubble extends only for a small distance inboard from the tips. The size of the hysteresis loop and the Reynolds number at which hysteresis starts was found to be influenced by the aspect ratio. The momentum deficit method was used to validate the data obtained by the strain gauge method and there was adequate agreement between the values found through the two methods. From the measurements of pressure done around the airfoil contour one could determine both the location of the laminar separation bubble and the regions were flow is separated. The pressure taps themselves were found to influence measurements somewhat in certain regions of angle of attack and Reynolds number. In the future it would be beneficial to continue strain gauge measurements on this airfoil with flaps and control surfaces to determine their effect on the formation of the laminar separation bubble. Also measurements on other shapes would give more insight into the phenomena occurring here. The effects of turbulence and noise will have to be investigated in detail to determine what performance to expect from an actual aircraft. Finally detailed measurements on boundary layer stability and its effect on the occurrence of reattachment should be studied in detail to gain insight into the reasons for the presence of a hysteresis loop in stall at these Reynolds numbers. / Ph. D.

A thin-film gage was used to measure the fluctuating component of heat transfer from a cylinder placed in a steady crossflow at Reynolds numbers, based on cylinder diameter, of approximately 19,000 and 30,000. Further, a one-dimensional flow pulsation at 13 Hz was added to the mean flow at a Reynolds number of 19,000, and the case of natural shedding locked on to exactly one-half the driving frequency was studied. A Gardon gage was also used to measure the time-averaged heat transfer under the same conditions. Both gages were mounted flush with the cylinder surface. The thin-film gage was maintained at a constant temperature by a constant temperature anemometer unit. A temperature controller actively matched the surrounding cylinder surface to the gage temperature to maintain a constant temperature boundary condition. The frequency response of the thin-film gage system was approximately 60 Hz. Representative time records of the instantaneous heat flux and the local fluid velocity were obtained at different locations by rotating the cylinder through 180°. Correlations between these signals in the time and frequency domain were also measured. Phase relationships between the unsteady heat transfer fluctuations and the local velocity were then obtained over the entire cylinder. In the attached boundary layer region, the heat flux signal was sinusoidal at the shedding frequency for the steady cases and at the driving frequency for the pulsating case. In the wake, however, the fluctuations were less organized in all cases. Here the magnitude of the fluctuations were higher than in the boundary layer region with peak-to-peak fluctuating amplitudes greater than 50% of the mean heat flux levels (as measured by the thin-film gage) over most of the wake. The phase relationship of the signals was nearly constant in the boundary layer region but varied with angular position in the wake indicating the existence of different flow regions in the wake. The local time-averaged heat transfer results reflect a 24% increase in heat transfer in the wake due to pulsation. This is a result of higher fluid velocities in the outer flow at 0 = 90° and in the wake regions when pulsations were added to the flow. A simple analytical model, based on the concept of an impinging jet that was oscillated back-and-forth across the wake, was developed to predict the heat transfer fluctuations in the wake. A parametric study was performed to determine the effect of changing the jet parameters on the predicted heat transfer fluctuations. Time records of the heat transfer fluctuations were obtained around the cylinder based on this model. The fluctuating amplitudes and phases of heat transfer (relative to the velocity) predicted by the model agreed well with the experimental values for the steady flow case. The increase of the local time-averaged heat transfer in the wake region due to flow pulsation was also predicted. / Ph.D.

No description available.

The aerodynamic performance of low aspect ratio wings at low Reynolds numbers applicable to micro air vehicle design was studied in this thesis. There is an overall lack of data for this low Reynolds number range, particularly concerning details of local flow behavior along the span. Experiments were conducted to measure the local pressure distributions on a wing at various spanwise locations in a Reynolds number range 30000 < Re < 90000. The model wing consisted of numerous wing sections and had a rectangular planform with NACA0012 airfoil shape with aspect ratio of one. One wing section, with pressure ports at various chordwise locations, was placed at different spanwise locations on a wing to effectively obtain the local pressure information. Integration of the pressure distributions yielded the local lift coefficients. Comparison of the local lift distributions to optimal elliptic lift distribution was conducted. This comparison showed a sharply peaked lift distribution near the wing tip resulting in a drastic deviation from the equivalent elliptic lift distributions predicted by the finite wing theory. The local lift distributions were further analyzed to determine the total lift coefficients vs angle of attack curves, span efficiency factors and the induced drag coefficients. Measured span efficiency factors, which were lower than predictions of the elliptic wing theory, can be understood by studying deviations of measured lift from the elliptic lift distribution. We conclude that elliptic wing theory is not sufficient to predict these aerodynamic performance parameters. Overall, these local measurements provided a better understanding of the low Reynolds number aerodynamics of the low aspect ratio wings.

An adaptive wing, a zero mass ux ow control device for low speed airfoil separation control, is investigated both experimentally and computationally at low speeds. The adaptive mechanism in the wings provides variable camber that can be actuated across a range of frequencies and amplitudes. Piezoelectric actuators are housed in a NACA 4415 airfoil with a chord length of :203 m. The entire adaptive wing assembly is then wrapped under a layer of latex membrane to provide a exible and smooth upper surface pro le. Experimental diagnostics include ow visualization, particle image velocimetry, as well as lift and drag measurements. The numerical simulation uses a 2D incompressible CFD code based on a nite-volume structured formulation with a chimera overset grid for the purpose of parallel computing. In the current study, the dimensionless speed range examined is 2:5 104 Re 1:5 105, where particular focus is given to Re 7:5 104, where Re = U` . All experiments and simulations are conducted in the range of attack angles from 0 24 and between reduced frequency values from 0 f+ 1:09, where f+ = f` U1 . Both experimental and computational results show that the region of separation is reduced when the actuation is turned on, thus enhancing aerodynamic e ciency. The maximum coe cient of lift increases by 26% when the reduced frequency, f+, is approximately :2, where the ow control mechanism appears to be most e ective. Phase-locked PIV and CFD vorticity plots reveal that the downward motion of the surface actuation decelerates the boundary ow and increases surface pressure, resulting in the formation of a series of cross-stream vortices that provides uid entrainment towards the suction surface, hence reducing separation.